Actuation system for aircraft engine lubrication pump and related aircraft

Abstract

The present invention relates to an actuation system (4), comprising: a generating rotary electric machine (6), comprising a first stator (14), comprising at least one output stator winding (16); a driving rotary electric machine (8), comprising a second rotor (18) intended to actuate a lubrication pump (2), and a second stator (20) comprising at least one input stator winding (22); a switch array (11), electrically connected to at least one output stator winding and at least one input stator winding, for electrically connecting them together or disconnecting them from each other according to a state of the switch array; and a control device (10), configured to control the switch array according to a state of a control signal, so as to electrically connect or disconnect at least one output stator winding and at least one input stator winding.

Description

Technical Field

[0001] The invention relates to an actuation system for a lubrication pump of an aircraft engine.

[0002] The invention also relates to an aircraft having such an actuation system.

[0003] The invention has application in the aviation field, in particular in the actuation of a lubrication pump of an aircraft engine. Background Art

[0004] Conventionally, in aircraft engines, in particular those of the UHBR ("Ultra High Bypass Ratio") type, a lubrication pump (called the main lubrication pump) is used, the task of which is to deliver oil to the bearings of the high-pressure and low-pressure shafts, as well as to the engine's transmission. This lubrication pump is driven by the high-pressure shaft via the auxiliary gearbox (or AGB).

[0005] Such engines also include an auxiliary lubrication pump, which is driven by the engine's low-pressure shaft and is used to assist the main lubrication pump in the event that it is no longer able to perform its function properly. This situation occurs in particular during certain flight missions in which the high-pressure shaft rotates at low speed, or even at airspeed, and the main lubrication pump is no longer driven. This situation particularly includes when the engine is shut down on the ground or in flight, or during an auto-spin phase (also known as free-spin) during the engine fan shutdown.

[0006] However, this architecture is not without its drawbacks. For example, since the low-pressure shaft can rotate in both directions, particularly at ground level under the influence of wind, it is necessary to use a gear system with a freewheel to ensure unidirectional rotation of the auxiliary lubrication pump.

[0007] Furthermore, if a continuously running auxiliary lubrication pump is employed, it is necessary to resize the lubrication circuit to take into account the inclusion of said auxiliary lubrication pump. The implementation of an occasionally running auxiliary lubrication pump is also problematic as long as a disconnectable auxiliary lubrication pump must be provided, which increases the complexity of its implementation.

[0008] Furthermore, the positioning of the auxiliary lubrication pump in the nacelle is constrained by the drive kinematics.

[0009] In order to overcome these problems, it is proposed to drive the auxiliary lubrication pump not via the low-pressure shaft but via an electric motor coupled to said auxiliary lubrication pump and said auxiliary lubrication pump is powered by a generator while being controlled via dedicated electronic control elements including power electronics.

[0010] Nevertheless, the solutions known from the prior art are not entirely satisfactory.

[0011] In fact, during operation, electronic control elements are often subjected to temperature conditions that can cause damage to the power electronics they contain, which creates reliability problems. Moreover, the reinforcement or thermal insulation of such electronic control elements usually translates into an increase in their mass and volume, which is disadvantageous in the present case.

[0012] One object of the present invention is therefore to propose an actuation system for a lubrication pump, in particular an auxiliary lubrication pump, of an aircraft engine that is reliable and robust while having minimal mass and volume. Summary of the Invention

[0013] For this purpose, the subject of the invention is an actuation system of the type mentioned above, comprising a generator rotary electric machine, a drive rotary electric machine, a control device and a switch matrix,

[0014] The generator rotary electric machine comprises a first rotor intended to be mechanically coupled to a shaft of the engine forming a transmission shaft, and a first stator comprising at least one output stator winding,

[0015] the driving rotary electric machine comprises a second rotor intended to be mechanically coupled to the lubrication pump for its actuation, and a second stator comprising at least one input stator winding,

[0016] a switch matrix being electrically connected to the at least one output stator winding on the one hand and to the at least one input stator winding on the other hand, for electrically connecting or disconnecting the at least one output stator winding and the at least one input stator winding from each other depending on a state of the switch matrix,

[0017] The control device is configured to control the switch matrix to electrically connect or disconnect the at least one output stator winding and the at least one input stator winding to each other according to a state of the control signal.

[0018] In practice, this actuation system does not include electronic control components that could malfunction due to the adverse temperature and vibration conditions in the nacelle. Instead, the voltage transmitted by the generator unit has the waveform necessary for the operation of the drive unit, allowing the drive unit to be driven in rotation directly by the voltage at the generator unit's output, without requiring voltage-shaping electronics, such as an inverter, to be arranged between the two rotating machines. This results in an actuation system for the lubrication pump that is more reliable and robust than known actuation systems, while also having a smaller mass and volume.

[0019] The actuation system according to the invention has another particularly advantageous technical effect.

[0020] In practice, the requirement in terms of oil flow rate on a lubrication pump is proportional to the rotational speed of the engine's shaft.

[0021] However, the actuation system according to the present invention is a linear system, that is, the rotational speed of the first rotor of the driving rotary electric machine is proportional to the rotational speed of the second rotor of the generator rotary electric machine. In other words, in the actuation system according to the present invention, the voltage at the terminals of each winding of the generator rotary electric machine has an amplitude and frequency that is a direct function of the rotational speed of the first rotor (and therefore of the drive shaft). This voltage applied to each winding of the driving rotary electric machine is converted into a rotational speed of the second rotor that is proportional to the rotational speed of the first rotor, and thus into the rotational speed of the drive shaft.

[0022] The above leads to an additional advantage of the invention, namely that, depending on the transmission shaft to which the first rotor is coupled (preferably the low-pressure shaft or the shaft of the engine's fan), and optionally via at least one suitable reduction gear, the speed at which the lubrication pump is driven is always sufficient to meet the requirements in terms of the oil flow rate required according to the rotational speed of the transmission shaft (that is to say according to the engine speed). Therefore, electronic control elements are superfluous.

[0023] According to further advantageous aspects of the invention, the actuation system comprises one or more of the following features, taken alone or in all technically possible combinations:

[0024] - the control device is configured to control the switch matrix to connect the at least one output stator winding and the at least one input stator winding to each other according to the rotation direction of the first rotor of the generator unit, so that the rotation direction of the second rotor remains unchanged regardless of the direction in which the first rotor of the generator unit is driven in rotation;

[0025] - the control device comprises an analog detection element configured to transmit a rotation signal representing a direction of rotation of the first rotor of the generator unit;

[0026] - the control device is configured to receive a control signal representing a rotational speed of a rotating element of the engine (preferably a transmission shaft), the control device being further configured to control the switch matrix so as to electrically connect the at least one output stator winding and the at least one input stator winding to each other if the rotational speed value of the rotating element is less than or equal to a predetermined threshold value, and, preferably, to electrically disconnect the at least one output stator winding and the at least one input stator winding from each other if the rotational speed value of the rotating element is greater than the predetermined threshold value;

[0027] - the switch matrix comprises at least one electromechanical switch;

[0028] The generator rotating electrical machine and / or the drive rotating electrical machine is a synchronous machine or an asynchronous machine with permanent magnets.

[0029] Furthermore, the present invention relates to an aircraft with a lubrication pump and an actuation system as described above, the first rotor of the generator rotary electric machine being mechanically coupled to the shaft of the aircraft's engine, and the second rotor of the drive rotary electric machine being mechanically coupled to the lubrication pump of the engine for actuation of the lubrication pump. BRIEF DESCRIPTION OF THE DRAWINGS

[0030] The invention will be better understood from the following description, given by way of non-limiting example only, and made with reference to the accompanying drawings, in which:

[0031] Figure 1 is a diagrammatic representation of an actuation system according to the present invention. DETAILED DESCRIPTION

[0032] Figure 1 A lubrication pump 2 and an actuation system 4 according to the invention are shown.

[0033] The lubrication pump 2 is intended to supply oil to elements to be lubricated, such as bearings or transmissions, of an aircraft engine (not shown).For example, the lubrication pump 2 is an auxiliary lubrication pump of the engine.

[0034] The actuation system 4 is used to extract energy at a shaft of the engine, called the transmission shaft, and use the extracted energy to actuate the lubrication pump 2. For example, the transmission shaft is the low-pressure shaft of the engine.

[0035] More precisely, the actuation system 4 comprises a generator rotary electric machine 6 , called generator unit, and a drive rotary electric machine 8 , called drive unit, which are advantageously electrically connected to one another via a switch matrix 11 controlled by a control device 10 .

[0036] The generator unit 6 is configured to absorb mechanical energy at the transmission shaft and convert the absorbed mechanical energy into electrical energy.

[0037] The drive unit 8 is configured to receive the electrical energy generated by the generator unit 6 and convert the received electrical energy into mechanical energy intended to actuate the lubrication pump 2 .

[0038] Preferably, each of the generator unit 6 and the drive unit 8 is a synchronous machine or an asynchronous machine having permanent magnets. In particular, the generator unit 6 is a synchronous machine or an asynchronous machine having permanent magnets, and the drive unit 8 is an asynchronous machine or a brushless DC motor. For example, the generator unit 6 is a synchronous machine having permanent magnets, and the drive unit 8 is an asynchronous machine.

[0039] The generator unit 6 includes a rotor 12 (referred to as a first rotor) and a stator 14 (referred to as a first stator).

[0040] The first rotor 12 is mechanically coupled to the drive shaft via any suitable mechanical transmission element, such as a reduction gear.

[0041] The first stator 14 comprises at least one winding 16 , referred to as an output stator winding, for example three output stator windings 16 .

[0042] The drive unit 8 includes a rotor 18 (referred to as a second rotor) and a stator 20 (referred to as a second stator).

[0043] The second rotor 18 is mechanically coupled to the lubrication pump 2 , directly or via at least one reduction gear, in order to cause its actuation.

[0044] The second stator 20 comprises at least one winding 22 , referred to as an input stator winding, for example three input stator windings 22 .

[0045] At least one output stator winding 16 and at least one input stator winding 22 are electrically connected to each other. In other words, each output stator winding 16 is connected to a corresponding input stator winding 22 without any power electronic components, such as transistors or thyristors, being arranged therebetween. Advantageously, each output stator winding 16 is connected to a corresponding input stator winding 22 via a switch of the switch matrix 11.

[0046] In this way, when the first rotor 12 of the generator unit 6 is driven in rotation, it circulates current between the at least one output stator winding 16 and the at least one input stator winding 22. This current causes the rotation of the second rotor 18 of the drive unit 8, which activates the lubrication pump 2.

[0047] Preferably, the control device 10 is configured to control the switch matrix 11 to electrically connect or disconnect the at least one output stator winding 16 and the at least one input stator winding 22 to each other. In particular, the control device 10 is configured to receive a control signal, for example, from a computer of the aircraft, and control the switch matrix 11 to connect or disconnect the at least one output stator winding 16 and the at least one input stator winding 22 to each other according to a state of the control signal.

[0048] According to another embodiment, the control signal is a rotational speed measurement signal transmitted by a sensor of the aircraft. This measurement signal represents the rotational speed of a rotating element of the engine, preferably the rotational speed of the engine's low-pressure shaft or fan. In this case, the control device 10 is configured to control the switch matrix 11 so as to electrically connect the at least one output stator winding 16 and the at least one input stator winding 22 to each other if the rotational speed of the rotating element is less than or equal to a predetermined threshold value. This predetermined threshold value may correspond to the occurrence of a situation in which at least one pump of the main lubrication system no longer rotates fast enough to properly ensure lubrication of the engine's transmission, necessitating the lubrication pump 2 to compensate for this shortfall as a pump of the transmission's auxiliary lubrication system. Preferably, the control device 10 is also configured to control the switch matrix 11 so as to electrically disconnect the at least one output stator winding 16 and the at least one input stator winding 22 from each other if the rotational speed of the rotating element is greater than the predetermined threshold value.

[0049] In this way, in the case where the lubrication pump 2 is an auxiliary lubrication pump, the lubrication pump 2 is advantageously activated only when the main lubrication pump needs assistance, in particular during an in-flight engine shutdown or during an automatic rotation phase of the engine's fan on the ground.

[0050] For example, the switch matrix 11 comprises a set of switches, preferably electromechanical switches, electrically connected to at least one output stator winding 16 on the one hand and to at least one input stator winding 22 on the other hand, the on or off state of each switch being controlled by the control device 10 .

[0051] Preferably, the control device 10 comprises an analog detection element 24 configured to transmit a rotation signal whose state is representative of the direction of rotation of the first rotor 12 of the generator unit 6 . This rotation signal also forms a control signal for the control device 10 .

[0052] For example, analog detection element 24 is configured to transmit a rotation signal based on a value derived from the voltage at the terminals of at least one output stator winding 16 of generator unit 6. In this case, control device 10 is further configured to control switch matrix 11 to induce phase rotation at drive unit 8, that is, to induce a change in the connection between at least one output stator winding 16 and at least one input stator winding 22, so that the rotational direction of second rotor 18 remains unchanged, regardless of the direction in which first rotor 12 of generator unit 6 is driven to rotate. In other words, regardless of the direction in which first rotor 12 of generator unit 6 is driven to rotate, second rotor 18 always rotates in the same predetermined direction, and therefore, lubrication pump 2 is always driven in the same direction.

[0053] The use of an analog detection element 24 is advantageous insofar as the robustness of the actuation system 4 is enhanced by avoiding the use of some or all digital components for transmitting the rotation signal.

Claims

1. An actuation system (4) for a lubrication pump (2) for an aircraft engine, the actuation system comprising a generator rotary electric machine (6), a drive rotary electric machine (8), a control device (10) and a switch matrix (11), The generator rotary electric machine (6) comprises a first rotor (12) intended to be mechanically coupled to a shaft of the engine forming a transmission shaft, and a first stator (14) comprising at least one output stator winding (16), The driving rotary electric machine (8) comprises a second rotor (18) intended to be mechanically coupled to the lubrication pump (2) for its actuation, and a second stator (20) comprising at least one input stator winding (22), The switch matrix (11) is electrically connected to at least one output stator winding (16) on the one hand and to at least one input stator winding (22) on the other hand, so as to electrically connect or disconnect the at least one output stator winding (16) and the at least one input stator winding (22) to each other according to the state of the switch matrix (11), The control device (10) is configured to control the switch matrix (11) to electrically connect or disconnect at least one output stator winding (16) and at least one input stator winding (22) to each other according to a state of a control signal, The control device (10) is further configured to control the switch matrix to connect at least one output stator winding (16) and at least one input stator winding (22) to each other according to the rotation direction of the first rotor (12) of the generator rotating electrical machine (6), so that the rotation direction of the second rotor (18) remains unchanged regardless of the direction in which the first rotor (12) of the generator rotating electrical machine (6) is driven to rotate.

2. The actuation system (4) according to claim 1, wherein The control device (10) comprises an analog detection element (24) configured to transmit a rotation signal representing the direction of rotation of a first rotor (12) of a generator rotary electric machine (6).

3. The actuation system (4) according to claim 1, wherein The control device (10) is configured to receive a control signal representing a rotational speed of a rotating element of an engine, and the control device (10) is further configured to control a switch matrix (11) so as to electrically connect at least one output stator winding (16) and at least one input stator winding (22) to each other if the rotational speed value of the rotating element is less than or equal to a predetermined threshold value.

4. The actuation system (4) according to claim 3, wherein: The rotating element of the engine is the drive shaft.

5. The actuation system (4) according to claim 3, wherein: The control device (10) is further configured to control the switch matrix (11) so as to electrically disconnect at least one output stator winding (16) and at least one input stator winding (22) from each other if the rotational speed value of the rotating element is greater than a predetermined threshold value.

6. The actuation system (4) according to claim 1, wherein The switch matrix (11) comprises at least one electromechanical switch.

7. The actuation system (4) according to claim 1, wherein: The generator rotating electrical machine (6) and / or the driving rotating electrical machine (8) are synchronous machines or asynchronous machines with permanent magnets.

8. An aircraft having a lubrication pump (2) and an actuation system (4) according to any one of claims 1 to 5, The first rotor (12) of the generator rotary electric machine (6) is mechanically coupled to the shaft of the aircraft's engine, The second rotor (18) of the driving rotary electric machine (8) is mechanically coupled to the lubrication pump (2) of the engine for actuation of the lubrication pump (2).

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